Automatic flight control system using instrument landing system information and including inertial filtering means for reducing ils noise

ABSTRACT

A flight control system includes localizer and glide path couplers using instrument landing system (ILS) information and inertially sensed data to generate appropriate control signals, and inertial filtering means for providing significant attenuation of ILS noise to reduce spurious signal generation.

Doniger et a1. Nov. 20, 1973 [54] AUTOMATIC FLIGHT CONTROL SYSTEM3,132,828 5 1964 Edinger et a1. 244 77 USING mSTRUMENT LANDING SYSTEM3,223,362 12/1965 Doniger 244/77 A 3,335,980 8/1967 Doniger et a1... 24477 A INFORMATION AND INCLUDKNG 3,447,765 6 1969 Doniger et al... 244 77A INERTIAL FILTERING MEANS FOR 3,489,378 1 1970 Watson et al.... 244 77A REDUCING 111s NOKSE 3,652,835 3/1972 Devlin et a1. 235 150.22

[75] Inventors: Jerry Doniger, Montvale; Frederic L. Swern, Bogota, bothof NJ.

Primary Examiner-Milton Buchler Assigneel The Bendix Corporation,Assistant ExaminerBarry L. Kelmachter Att0rneyAnth0ny F. Cuoco et a1.

[22] Filed: June 18, 1971 [21] Appl. No.: 154,517

[57] ABSTRACT [52] US. Cl. 244/77 A, 244/77 G [51] Int. Cl. B644: 13/50A flight control system includes localizer and glide Field of Search244/77 R, 77 A, 77 D, path couplers using instrument landing system(ILS) 244/77 G, 3.15, 3.19, 3.2, 3.21; 343/108 R, information andinertially sensed data to generate ap- 108 M, 107; 235/ 150.22; 318/489propriate control signals, and inertial filtering means for providingsignificant attenuation of ILS noise to [56] References Cited reducespurious signal generation.

UNITED STATES PATENTS 2,981,500 4/1961 Carlton 244/3.15 7 Claims, 4Drawing Figures K, INTEGRATOR P'TCH ERROR 40 l 32 MEANS S (2313121. 8 lI 1 com/mo PITCH 22 MODIFIER 212 I LAG c FILTER f- /(2) 4 T 24 D 0 p i/5 DE E M 6 l 42 MEANS IO 2 ig g LIMITER NORMAL LAG. ACCEL. FILTERPAIENTEUNUVZO ms 23.773; 281

WEI 3c; 3

I NTEGRATOR INERTIAL PLATFORM SYSTEM INTEGRATOR INERTIAL PLATFORM SYSTEMFIG. 4

INVENTOR5 JERRY DO/V/GEA FPEDER/C LSWEPN HTTOQNEY AUTOMATIC FLIGHTCONTROL SYSTEM USING INSTRUMENT LANDING SYSTEM INFORMATION AND INCLUDINGINERTIAL FILTERING MEANS FOR REDUCING ILS NOISE BACKGROUND OF THEINVENTION 1. Field of the Invention This invention relates to flightcontrol systems and, particularly, to flight control systems usinglocalizer and glide path couplers. More particularly, this inventionrelates to flight control systems of the type described including meansfor reducing the sensitivity of instrument landing system (ILS) approachcoupling systems while maintaining accurate control of the aircraftunder gust conditions with attendant low levels of spurious cockpitcontrol activity or aircraft motion.

DESCRIPTION OF THE PRIOR ART Most flight control systems with approachcoupling means use standard instrument landing systems (ILS) localizerand glide path radio guidance signals as references. The localizerguidance signal is referenced to the extended centerline of the runway.The glide slope guidance signal defines a vertical path that intersectsthe approach end of the runway at a prescribed angle from thehorizontal, generally about 2.7".

The radio guidance provides proportional present position angular datawith respect to the effective origin of the transmitted signals. Theseradio signals are affected by terrain variations including hills orlarge buildings, or over-flying aircraft that act as reflectors. Thetotal, direct and reflected, position data received in the aircraftprovides the useful long term track references as well as providingshort term noise.

This noise is a major problem associated with the design of ILS coupledautomatic and manual flight control systems. These systems must havehigh gain to force the aircraft to maintain an accurate approach trackthat will insure desirably small lateral and longitudinal landingdispersions, particularly under various wind conditions. The high gain,however, makes the approach coupler configurations sensitive to thereceived beam noise, particularly at frequencies that result in rapidcockpit control and display activity as well as spurious aircraftrolling and pitching activity. It is important therefore, that amechanization be provided to significantly reduce the sensitivity of ILScoupling systems while maintaining accurate control of the aircraftunder gust conditions with attendant low level spurious cockpit controlactivity or aircraft motion.

Flight path integrations due to aircraft motion force the systemdesigner to provide, in the ILS coupler mechanization, some sort ofdamping term to maintain desirable system stability. Generally, in orderto provide damping for a beam displacement control loop, some measure ofbeam rate is necessary. Explicit rate filters have been used in thepast. They provide stable performance and accurate tracking under windconditions, but at the unacceptable sacrifice of spurious controlsurface and attitude responses when subjected to the aforenoted ILSnoise.

Approximate measures of beam rate can be generated by'using changes inheading for localizer and changes in pitch attitude for glide path. Thishas been successfully used in approach couplers for many years, with theprimary advantage being that the measurement is derived from inertialreferences which already exist for use in the autopilot or flightdirector. The disadvantages are that path tracking is inaccurate undergust or wind conditions and the beam displacement signals areunfiltered.

In order to improve tracking accuracy under wind conditions, the beamrate parameter must be increased and the system noise responses thusdegraded. The wind performance is degraded by systems using thisprinciple since gusts require the aircraft to weathercock to minimizelateral or normal accelerations which move the aircraft away from itsintended track. Explicit heading and pitch attitude control inhibitsthis desirable gust weathercocking, and the aircraft is forced todeviate from the intended path.

In order to overcome these disadvantages, localizer coupler arrangementshave used a lagged function of roll attitude to provide primary pathdamping and veinier low gain damping is provided by beam rate andheading. The lagged roll parameter approximates the integral of rollthat is proportional to the change in heading, or beam rate. This termis effective since it does not interfere with the lateral weathercockingof the aircraft in gusts.

However, systems which depend on a lagged roll function alone fordamping require sufficiently high gain to be susceptible to lowamplitude, long period oscillations that result from the lag filter notbeing a pure integration. A pure integration cannot be used sincepractical system design must allow for drift of integration devices. Inaddition, a separate integrator is normally used to provide trimcompensation and the two integrators provide an undesirable oscillatorymode for the system. Therefore, beam rate and/or heading terms are usedin conjunction with the lagged roll term. These added terms, however,must be carefully tailored to minimize beam noise effects and wind gustperformance degradation.

A filtered altitude rate signal has been used in the pitch axis toimprove wind gust performance of the attitude reference (glide slope)system. The altitude rate loop is stabilized by a rapidly washed outbarometric altitude signal and a pitch rate Signal, and which signalsare not significantly involved with the aircraft track responses. Theundesirable feature of the altitude signal is its susceptibility to gustnoise and local flow effects at the static ports which introduceadditional noise into the system. In addition, the need for barometricrefer ence during a final approach, particularly in redundant systems,requires considerable system complexity. Also, for large aircraft,ground effects create additional disturbances on the altitude ratesignal at low altitudes.

A modified glide path coupler arrangement in which the lagged normalacceleration term is substituted for the barometric altitude rate signalhas been used. This arrangement provides improved wind shear performanceover the attitude system without the attendant beam noise sensitivity ofthe rate system. However, the use of a simple body mounted accelerometeris subject to null offsets which are propogated through the high gainlag filter to drive the aircraft from the ILS glide path reference.

SUMMARY OF THE INVENTION The flight control system of the inventionprovides excellent wind performance (track accuracy) and low noisesusceptibility, as compared with the pure lagged roll system forlocalizer or the attitude damped system for glide path, and very lowsusceptibility to the long term tracking inaccuracies that result fromroll attitude or normal acceleration offsets. Accordingly, the inventioncontemplates providing a heavily filtered rate signal which is formed toprovide a good low frequency beam rate reference. High frequency noise(nominally have 0.1 radian per second) is increasingly attenuated. Ashort term beam rate reference signal is provided using filtered rollattitude in localizer and filtered normal acceleration in glide path. Inthis manner low frequency efiects such as null offsets of the inertialsensors are greatly attenuated. The low and high frequency beam rateterms are combined to form a single wide band, relatively noise freebeam rate signal. The beam rate signal is limited to further restrictthe noise response of the system and inertial navigation apparatus canbe used to provide further noise immunity.

One object of this invention is to provide a flight control system usinglocalizer and glide path couplers responsive to ILS information andinertially sensed data to generate appropriate control signals, andincluding means for providing significant attenuation of ILS noise toreduce spurious signal generation.

Another object of this invention is to significantly reduce thesensitivity of ILS approach coupling systems, while maintaining accuratecontrol of the craft under gust conditions, with attendant low levels ofspurious cockpit control activity or aircraft motion.

Another object of this invention is to provide a flight control systemhaving excellent wind gust performance (track accuracy), low noisesusceptibility and low susceptibility to long term tracking inaccuraciesthat result from roll attitude and normal acceleration offsets.

Another object of this invention is to use the inertial elementsnormally existing in flight control systems to provide significantfiltering of ILS localizer and glide path beam noise while providingaccurate well-damped performance.

Another object of this invention is to utilize inertial navigationsystem data to provide further noise immunity. 7

Another object of this invention is to provide strategically located ILSbeam rate signal limits for further reducing large amplitude noiseattenuation of the systern.

The foregoing and other objects and advantages of the invention willappear more fully hereinafter from a consideration of the detaileddescription which follows, taken together with the accompanying drawingswherein several embodiments of the invention are illustrated by way ofexample. It is to be expressly understood, however, that the drawingsare for illustration purposes only and are not to be construed asdefining the limits of the invention.

DESCRIPTION OF THE DRAWINGS FIG. 1 is a block diagram of a flightcontrol system including a glide slope coupler according to theinvention.

FIG. 2 is a block diagram of a flight control system including alocalizer coupler according to the invention.

FIG. 3 shows an embodiment of the device of FIG. I using inertialnavigation system (INS) data.

FIG. 4 is an embodiment of the device of the invention shown in FIG. 2using inertial navigational system (INS) data.

DESCRIPTION OF THE INVENTION With reference to FIG. 1, a beamdisplacement signal means 2 provides a signal corresponding to thedisplacement of an aircraft from a predetermined glide slope reference.The beam displacement signal is applied to a desensitizer 4 whichdesensitizes the signal as a function of radar altitude and provides asignal nearly proportional to the change in aircraft altitude A h. Thedesensitized beam displacement signal (A h) is applied to a-summingmeans 6 and to a rate filter 8 having a time constant T The filteredsignal is applied to a limiter 10 which provides an altitude rate signalA normal accelerometer 12 provides a normal acceleration signal A whichis applied to a lag filter 14 having a time constant T for providing asignal Ii nearly proportional to altitude rate. Signal Ii is applied toa summing means 16 and summed thereby with the output of an integrator19, which integrates the signal from a summing means 18 to provide asignal 11 Signal H from summing means 16 is fed back to a summing means18 andsummed thereby with signal hi from limiter 10, and the summedsignal is integrated by integra tor 19. Integrator 19 has a timeconstant T l/S.

Signal h from summing means 16 is applied through an amplifier 20 havinga gain k to summing means 6 and summed thereby with signal A It fromdesensitizer 4. The summation signal from summing means 6 is applied toa lag filter 22 having a time constant T for providing a pitch commandsignal 0 Signal h from summing means 16 and signal 6 from lag filter 22are applied to a summing means 24 and the summation signal is applied toa command modifier 26. Signal A h from desensitizer 4 is applied throughan amplifier 28 having a gain K,to an integrator 30. the signal fromintegrator 30 is applied to a summing means 32 and summed thereby withthe signal from command modifier 26. The signal from summing means 32 isapplied to a summing means 34 and summed thereby with a signal from apitch error means 36 and a signal from a pitch rate means 38. Summingmeans 34 provides an elevator control signal 8 which is applied to acontrol means 40, which may be a conventional servo system, forcontrolling aircraft elevators 42.

Signal h from summing means 16 is a linear combination of filtered beamdisplacement and acceleration signals and may be expressed as follows:

h" h (TAS/TA 5+1 Hi, my 3+1 where ii A, T /I S+l (2) where An/S Ii atfrequencies above T,,] radians per second and where;

A [1 SIT S-l-l where Ii, Ii at frequencies below T I radians per second.If T l T l and T 1 T 1, then equation 1 provides that signal H isproportional to A ii about the glide path. This signal is also used todrive elevator 42 as a primary path damping term and is used tocomplement the filtered beam displacement terms as follows:

if h h' under noise free conditions and 5 li=sAh, then 9;=1AhKs+1/T,,s+1 I ifK T then K A h With reference to FIG. 3, it will beseen that signal H can be derived from an inertial platform system 44.Thus, signal ii from inertial platform system 44 is combined with thesignal from integrator 18 by summing means 16 to provide signal ii.Since inertial platform systems operate on vertical acceleration, theuncertainties relative to the use of a body mounted normalaccelerometer, such as normal accelerometer 12, are much smaller, andfilter time constants T and T, can be increased to provide further ILSnoise rejection.

With reference, now, to FIG. 2, a beam displacement signal means 50provides a signal corresponding to the displacement of the aircraft froma predetermined localizer reference. The signal from beam displacementmeans 50 is applied to a desensitizer 52 which provides a signal A Ywhich is approximately proportional to a cross track reference. Signal AY is applied to a summing means 54 and to a rate filter 56 having a timeconstant T The signal from rate filter 56 is applied to a limiter 58which provides a cross track rate signal Y,, and which signal Y, isapplied to a summing means 60.

A signal from a roll altitude sensor 62 is applied to a lag filter 64having a time constant T, for providing a cross track rate signal jaSignal y is applied to a summing means 66 and summed thereby with theoutput of an integrator 68 having a time constant T l /S to provide asignal y.

Signal y from summing means 66 is fed back to summing means 60 andsummed thereby with signal Y from limiter 58. The summation signal fromsummation means 60 is integrated by integrator 68 and the integratedsignal is applied to summing means 66 and summed with the signal fromfilter 64 to provide signal Signal is applied through an amplifier 67having a gain K to summation means 54 where the amplified signal issummed with signal A Y from desensitizer 52. The summation signal fromsummation means 54 is applied to a lag filter 70 having a time constantT and therefrom to an amplifier 72 having a gain K Amplifier 72 providesa roll command signal 42 Signal qb from amplifier 72 is applied to asummation means 74 and summed thereby with signal Y' from summationmeans 66. The summation signal from summation means 74 is applied to acommand modifier 76. The signal from command modifier 76 is applied to asummation means 78 and summed thereby with signal A Y from desensitizer52 applied through an amplifier 80 having a gain K and through anintegrator 82.

The signal from summation means 78 is applied to a summation means 84and summed thereby with a signal from a roll rate sensor 86 and with asignal from a roll error sensor 88. The summation signal from summationmeans 84 is a roll control signal 8 which is applied to a control means90, which may be a conventional type servo system, for controllingailerons 92 of the aircraft.

As shown in FIG. 2, signal is related to signal A y and roll attitude asfollows:

ms/1.9+ 1) 'nu/TAs 1 /3 The aircraft cross track rate that results whenthe bank angle, qYchan ges is y gxb/s, where g= acceleration to duegravity. At frequencies above T 1 radians per second, lag filter 64operates as an integrator and its output relative to (b is proportionalto y. Signal y follows signal y, for frequencies above TA 1 radians persecond and attenuates the signal for lower frequencies. The term is across track rate derived by a lead network as follows:

y A y (S/T S +1),

where T is very small relative to T At frequencies below T 1 radians persecond, beam rate signal y is passed and at higher frequencies, the beamrate term is attenuated:

Since, under noisefree conditions, y; then c= D( D U+ y 1 =K y (K S l/TS +1) Setting K T causes the output of filter to be a wideband versionof the true aircraft track deviation. Beam noise is attenuated forfrequencies above T 1 radians per second.

FIC shows that the filtered roll attitude sensor which generates theshort term cross track signal Y can be replaced by an inertial platformsystem 92.

Thus platform system 92 generates signal Y, and which signal is combinedby summation means 66 with the signal from integrator 68. The same kindof long term uncertainties exist for this signal, although at a lowerlevel as for the roll offsets. Therefore the filter time constants T andT can be increased, when the INS Y reference is used, to further reducethe effects of [LS noise.

From the aforegoing description of the invention with reference to thedrawings, it will be seen that both the localizer and glide path systems are based on an arrangement whereby heavily filtered rate signals areformed to provide a good low frequency beam rate reference. Highfrequency noise (nominally above .1 rad/- see. is increasinglyattenuated. A short term beam rate reference signal is formed usingfiltered roll attitude in localizer and filtered normal acceleration inglide path. In this manner, low frequency effects such as null offsetsof the inertial sensors are greatly attenuated.

The low and high frequency beam rate terms are combined to form a singlewideband, relatively noise free, beam rate signal. This method offorming a useful signal from two independent sources is called thecomplementary filter method. The complemented beam rate signals are usedin two ways; they provide primary damping for the ILS control loops andthey are used in a second stage of complementation to filter the beamdisplacement signal to further reduce the beam noise susceptibility ofthe system.

It will be understood that the components of the invention describedherein are all of the conventional type well known in the flight controlart. The novelty of the invention resides not in the componentsthemselves but in their arrangement as shown in the drawmgs.

Although several embodiments of the invention have been illustrated anddescribed in detail, it is to be expressly understood that the inventionis not limited thereto. Various changes may also be made in the designand arrangement of the parts without departing from the spirit and scopeof the invention as the same will now be understood by those skilled inthe art.

What is claimed is:

1. An aircraft control system comprising:

means for providing a signal corresponding to the displacement of thecraft from a predetermined reference beam;

means for providing a low frequency beam rate signal;

means for providing a high frequency beam rate signal;

means connected to both of the beam rate signal means and to the beamdisplacement signalmeans and responsive to the signals therefrom forproviding a control signal including means for combining the high andlow frequency beam rate signals, means for desensitizing the beamdisplacement signal, means for combining the desensitized beamdisplacement signal and the combined beam rate signal, means forfiltering the combined signal from the last mentioned means to provide acommand signal and means connected to the command signal means forproviding the control signal in response to the signals therefrom; and

means connected to the control signal means and responsive to thecontrol signal for controlling the craft.

2. An aircraft control system as described by claim 1, wherein the meansfor providing a high frequency beam rate signal includes:

means for providing a normal acceleration signal;

and

means having a predetermined lag constant for filtering the normalacceleration signal to provide the high frequency beam rate signal, saidsignal being nearly proportional to aircraft altitude rate.

3. An aircraft control system as described by claim 1, wherein the meansfor providing a low frequency beam rate signal includes;

means for desensitizing the beam displacement signal;

means for filtering the desensitized signal;

means for limiting the filtered signal; and

means for integrating the limited signal.

4. An aircraft control system as described by claim 3, including:

means for combining the integrated signal and the high frequency beamrate signal; .means for combining the combined signal and the limitedsignal; and

the integrator integrating the signal from the last mentioned combiningmeans.

5. An aircraft control system as described by claim 1, wherein the meansconnected to the desensitizing means and to the command signal means forproviding the control signal in response to the signals therefromincludes:

means for integrating the desensitized signal;

means for modifying the command signal;

means for combining the integrated signal and the modified signal;

means for providing an attitude error signal;

means for providing an attitude rate signal; and

means for combining the combined integrated and modified signal, theattitude error signal and the attitude rate signal to provide thecontrol signal.

6. An aircraft control system as described by claim 5, wherein:

the means for providing an attitude error signal provides a pitch errorsignal; and

the means for providing an attitude rate signal provides a pitch ratesignal.

7. An aircraft control system as described by claim 5, wherein:

the means for providing an attitude error signal provides a roll errorsignal; and

the means for providing an attitude rate signal pro-

1. An aircraft control system comprising: means for providing a signalcorresponding to the displacement of the craft from a predeterminedreference beam; means for providing a low frequency beam rate signal;means for providing a high frequency beam rate signal; means connectedto both of the beam rate signal means and to the beam displacementsignal means and responsive to the signals therefrom for providing acontrol signal including means for combining the high and low frequencybeam rate signals, means for desensitizing the beam displacement signal,means for combining the desensitized beam displacement signal and thecombined beam rate signal, means for filtering the combined signal fromthe last mentioned means to provide a command signal and means connectedto the command signal means for providing the control signal in responseto the signals therefrom; and means connected to the control signalmeans and responsive to the control signal for controlling the craft. 2.An aircraft control system as described by claim 1, wherein the meansfor providing a high frequency beam rate signal includes: means forproviding a normal acceleration signal; and means having a predeterminedlag constant for filtering the normal acceleration signal to provide thehigh frequency beam rate signal, said signal being nearly proportionalto aircraft altitude rate.
 3. An aircraft control system as described byclaim 1, wherein the means for providing a low frequency beam ratesignal includes; means for desensitizing the beam displacement signal;means for filtering the desensitized signal; means for limiting thefiltered signal; and means for integrating the limited signal.
 4. Anaircraft control system as described by claim 3, including: means forcombining the integrated signal and the high frequency beam rate signal;means for combining the combined signal and the limited signal; and theintegrator integrating the signal from the last mentioned combiningmeans.
 5. An aircraft control system as described by claim 1, whereinthe means connected to the desensitizing means and to the command signalmeans for providing the control signal in response to the signalstherefrom includes: means for integrating the desensitized signal; meansfor modifying the command signal; means for combining the integratedsignal and the modified signal; means for providing an attitude errorsignal; means for providing an attitude rate signal; and means forcombining the combined integrated and modified signal, the attitudeerror signal and the attitude rate signal to provide the control signal.6. An aircraft control system as described by claim 5, wherein: themeans for providing an attitude error signal provides a pitch errorsignal; and the means for providing an attitude rate signal provides apitch rate signal.
 7. An aircraft control system as described by claim5, wherein: the means for providing an attitude error signal provides aroll error signal; and the means for providing an attitude rate signalprovides a roll rate signal.